Missile directional control system



Feb. 14, 1967 LUDTKE 3,304,029

MISSILE DIRECTIONAL CONTROL SYSTEM Filed Dec. 20, 1963 2 Sheets-Sheet lINVENTOR. Nor 27747? 7. 1 21/2 J0 ,9 rTaF/VI/J Feb. 14, 1967 N. F.LUDTKE MISSILE DIRECTIONAL CONTROL SYSTEM 2 Sheets-Sheet 2 Filed Dec.20, 1963 1 M 4 O 4 \fl w \7Y M w 3% H w/ l Z} n J, C 1 1 1; Q? q \1 w 7J J 4 4 i j 4 J i J #4 4 M iTrax/vz'ng United States Patent 3,304,029MISSILE DIRECTIONAL CONTROL SYSTEM Norman F. Ludtke, WashingtonTownship, Macomb County, Mich., assignor to Chrysler Corporation,Highland Park, Mich., a corporation of Delaware Filed Dec. 20, 1963,Ser. No. 332,127 18 Claims. (Cl. 244-320) This invention relatesgenerally to missiles and, more particularly, to an improved directionalcontrol system therefor.

The directional control of a missile or similar vehicle is accomplishedby a system that balances the forces acting on the missile normal to theflight path during the boost phase of flight. These forces are thrust,drag, aerodynamic normal forces, and control forces. If these forces arein balance, ignoring gravity, the missile will fly a straight line to atarget comprising any object or point in space.

Directional control systems for missiles and similar type vehiclescurrently known and in use employ a great variety of relatively complexelectronic, mechanical, and pneumatic systems to accomplish the guidanceand control function. Such systems are not justified in relatively shortrange missiles, particularly of the small artillery type. Prior tolaunching, such a missile is aimed with its longitudinal axis accuratelyaligned with its target or along its predetermined path. Afterlaunching, the missile is subjected to wind and other disturbing forcesnormal to its predetermined path causing the missile to pitch or yawabout its center of gravity and to deviate from the desired attitude. Itis necessary to sense the occurrence of this deviation and to provide abalancing force promptly to maintain the missile on its predeterminedpath.

It is an object of this invention to provide a directional controlsystem for a missile in which a single gyroscope element is utilizedboth to sense and to directly initiate and control the balancing forcesrequired.

It is a further object of this invention to provide a directionalcontrol system for a missile in which a single energy source is utilizedto provide gyroscope spin and to provide and maintain a constantpressure source for initiating and maintaining corrective forceapplication.

It is a still further object of this invention to provide an improvedmissile directional control system in which the deviating movement ofthe missile itself from its launch predetermined path serves to generatea counterbalancing force by corrective jets without the interveningnecessity for transducers.

The foregoing objects and others will become aparent from the followingdiscussion and description of illustrative forms of my invention. In thedrawings ac companying the description, in which reference charactershave been used to designate like parts referred to in the description:

FIGURE 1 is a view of a missile incorporating the present invention withparts broken away and in partial cross section to better illustrate itsplacement and mode of operation;

FIGURE 2 is a sectional view to greatly enlarged scale of a portion A ofFIGURE 1;

FIGURE 3 is a section taken along the line BB of FIGURE 2;

FIGURE 4 is a view similar to FIGURE 2 showing an alternate embodimentof the present invention; and

FIGURE 5 is a section taken along the line C-C of FIGURE 4.

The invention is embodied in the missile illustrated in FIGURE 1 of thedrawings. The numeral designates the motor for the missile with motornozzle 12. A suitable fuel for the rocket motor is indicated in chamber14 forward of the motor. A Warhead 16 is shown mount ed in the missile.Fins 18 are mounted on the aft portion of the missile to provideaerodynamic stability in flight. The directional control system whichforms the subject matter of the present invention is shown with itsoperating elements mounted about the longitudinal axis 20 of the missileand around rocket motor nozzle 12. The directional control systemalternately may be positioned in a like manner but forwardly of themissile center of gravity. A plurality of nozzles 22 are bodily fixed tothe missile and oriented to provide corrective jet forces thereto tomaintain the missile on its predetermined path in accordance with thepresent invention. A two degree of freedom gyro 24 is mounted on themissile with its spin axis coincident with the axis of the missile.Stator 26 is mounted about the motor nozzle 12 as shown in FIG. 2. Rotor28 is mounted for rotation about stator 26 on a double ball bearingraceway 30. Included in raceway 30 are balls 32, an inner race 34attached to stator 26, an outer race 36 attached to rotor 28, and a snapring 38 for retaining the balls in place. A solid propellant 40 iscontained in rotor 28 for rotating it at a desired rate. To that end, aplurality of spin orifices 42 arranged in two spaced and concentricrings are provided through the rotor 28. Spin orifices 42 are aligned toprovide tangential forces to provide a balanced rotation of rotor 28responsive to burning of the propellant 40. It will further be seen thatpressure emitted through spin orifices 42 is communicated to plenumchamber 48 to provide a source of operating pressure therefor. Rotor 28further is provided on its outer periphery with a peripheral groove 44.positionable in correspondence with the inlet portion 46 of each of thenozzles 22. In the practice of the present invention, at least threenozzles 22 are required, equally spaced about the missile. In thepreferred embodi' ment here employed, four nozzles are utilized, two inthe pitch plane and two in the yaw plane of the missile. Plenum chamber48 is bodily fixed to the missile and mounted on stator 26 enclosingrotor 28 with a chamber adapted to be pressurized to provide fluid forcorrective force jets. Plenum chamber 48 further acts as a valveoperating means by reason of the operation of its nozzle inlet portions46 in cooperation with grooved portion 44 of rotor 28 in a manner to beexplained in the section Description of Operation" hereinafter.

FIGURE 2 is further illustrative of the detail of the directionalcontrol system elements. A number of caging systems could be utilized.In the present system an inertial caging system 50 is incorporated whichincludes spring loaded plunger 52, locking pin 54, and locking groove56. The gyro is shown in its caged, pre-launch position. Solidpropellant 40 comprises a propellant of the cool burning type i.e. of atype combustible at about 2,000" F. A layer of readily ignitablematerial 58 is utilized to promote uniform and instantaneous ignition ofpropellant 40. To initiate the firing of the propellant, a source ofelectrical potential 60 is utilized. An extremely light and frangiblegrounded conductor 62 is threaded through nozzle 22, inlet portion 46,groove 44 and one of the spin orifices 42 to provide an electricalimpulse to fire the solid propellant. The interior of plenum chamber 48is further provided with a plurality of baflle plates 64 which arecircumferentially and equally spaced about the chamber intermediateinlet portions 46 in the manner shown. It is the function of batheplates 64 to reduce the fluid velocity in the plenum chamber in the areaof the inlet portions 46. This serves to promote more uniform and lessturbulent flow from the nozzles 22 and to minimize undesirable torque onthe rotor 28 near its groove portion 44.

FIGURE 3 shows a section taken along the line BB in FIGURE 2 andillustrates the cross-sectional configuration of inlet portions 46.These inlet portions are of a rectangular cross-section to promotelinearity of operation as the inlet portions 46 carried by plenumchamber 48 are rotated about the groove portion 44 of rotor 28.

FIGURE 4 illustrates an alternate embodiment of the present invention inwhich an external source of pressurized fluid 66 is utilized both topressurize plenum chamber 48 and to provide a rotative torque to spinrotor 28 up to the required angular rate. Valves 68 and 70 are providedin the supply lines from source 66 to provide the required fiow rates inconduits 72 and 74, respectively, to achieve these purposes. Conduit 72extends through the wall of plenum chamber 48 through its upper end. Abafiie 73 is mounted across the outlet from conduit 72 to prevent directimpingement of the gas on rotor 28. Conduit 74 extends through the lowerend of plenum chamber 48 and terminates in an outlet nozzle 75 fixed tothe inner surface of chamber 48. Nozzle 75 is directed to apply a fluidjet against a series of notches 45 formed in the bottom of groove 44 asbest shown in FIGURE 5. Each notch 45 extends across the bottom ofgroove 44 and is so formed that the impingement of the jet from nozzle75 rotates rotor 28 about its spin axis in a manner well known in theart.

Description of operation After acquisition of a target, the missile ispositioned with its axis 20 in alignment with the target by alineof-sight aiming procedure. With reference to FIGURE 2, an electricalimpulse from source 60 is furnished through lead 62 to ignite igniterlayer 58 and solid propellant 40. It should be noted that propellant 40is of a sufficient quantity to burn for the entire length of the missileflight. Responsive to the burning of propellant 40, gas is expelled fromthe two rows of spin orifices 42. The center line of each orifice is ina plane that passes through the center of motion of the gyro. By settingthe center line of each orifice at an angle so that it passes the centerof motion at a predetermined distance, the required rotative torque isapplied to spin the rotor 28. As soon as rotation of rotor 28 begins,frangible lead 62 is broken loose and is spun loose by the rotation ofrotor 28. To obtain the desired rpm. the propellant is ignited prior tothe launch of the missile. It will be seen that by suitable alterationof the geometry of the burning surface of propellant 40, adjustment maybe made of the length of the spin-up time. During spin-up time, the gaspressure from orifices 42 is communicated to plenum chamber 48 therebypressurizing the chamber. In

the caged condition of the gyro prior to launch, the rotor 28 ismaintained by inertial caging system 50 in the position illustrated inFIGURE 2. In the caged position of rotor 28, its groove 44 is positionedwith its lower edge portion centered in the plurality of inlet portions46 formed in plenum chamber 48. Thus, one half of each nozzle inletportion 46 is covered by the rotor 28 and the other half is exposed bygroove 44. Accordingly, the fiow supplied in the caged condition is thesame for each nozzle inlet portion 46 and its corresponding jet nozzle22. The width of the groove 44 is further sized so that at the expectedmaximum angular rotation of the missile the throat area uncovered of aninlet portion 46 in that direction is fully open.

When sufficient time has elapsed to spin the gyro at the desired rateand to pressurize plenum chamber 48, the missile is launched by firingrocket motor 10. Inertial caging system 50 operates to move plunger 52rearwardly to lock plunger 52 in groove 56 whereby relative movementbetween plenum chamber 48 and rotor 28 is allowed and the directionalcontrol system becomes operative. Plenum chamber 48 is bodily fixed androtates with the missile. During flight the missile is caused to changeits attitude from its predetermined line of flight by aerodynamicdisturbance. This motion serves to decrease or increase the throat areaprovided by inlet portions 46 and consequently decrease or increase thejet forces applied through jet nozzles 22. With four equally spacednozzles, the motion of the missile caused by the action of plenumchamber 48 increases in How from one jet nozzle 22 and decreases thefiow from the opposite nozzle. The difference between the two forcesapplied results in a net corrective force in the required direction toprovide a balance of forces that will maintain the missile on itspredetermined'path. By proper sizing of the nozzles and adjustment ofplenum chamber pressure, the required control forces for directionalcontrol of the missile may be achieved. It should be noted that the endchambers of the plenum chamber are sized so that they extend well beyondthe edges of the grooves 44 on both sides. The air gap between thenozzle inlet portions 46' and the peripheral surface of rotor 28 is heldto a minimum. This serves to minimize gas flow into the nozzles normalto the direction of spin. Inlet portions 46 further protrude into plenumchamber individually and are not part of a concentric ring. This permitthe pressurized gas to feed into groove 44 between nozzle inlet portions46. Gas flow through nozzle inlet portions 46 is permitted only in thedirection parallel to rotor spin. This serves to eliminate torque inrotor 28 due to cross flows through the throats of nozzle inlet portions46. The system is characterized by a self filtering feature due to theexpulsion of gas by spinning rotor 28 through orifices 42. This motionimparts a centrifugal force on the gas thereby forcing the heavierparticles to the outer face of the plenum chamber.

FIGURES 4 and 5 illustrate an alternate embodiment of the presentinvention in which an external source of constant pressure 66 isselectively connected both to provide rotor spin and to provide therequired operating pressure for plenum chamber 48. A further embodimentof the present invention might incorporate a solid propellant after themanner of FIGURE 2 to provide gyro spin and an external pressure sourceapplied after the manner of FIGURE 4 to supplement in the pressurizingof plenum chamber 48.

It will be seen that I have provided a directional control system whichis characterized by its simplicity and improved mode of operation andreadily applicable to a great variety of missiles or, indeed, to anybody thrust propelled in a medium that allows two degrees of freedommotion.

I claim:

1. A directional control system for maintaining a missile on a launchpredetermined path comprising a gyroscope mounted with its spin axis androtor aligned with the longitudinal axis of the missile at launch, saidmissile movable transaxially relative to said rotor, at least threesubstantially equally spaced and radially aligned jet nozzles mounted onthe missile, a source of pressurized fluid operatively connectible tosaid nozzles, said gyroscope having said rotor operatively connectedbetween said source and said nozzles for providing a variable fluid flowtherethrough to provide balancing forces to the missile responsive toits movement relative to said rotor.

2. The combination as set forth in claim 1 in which said source ofpressurized fluid comprises a plenum chamber fixed to said missile,mounted about and enclosing said rotor and operatlvely connected to asource of positive pressure, said chamber defining a plurality ofvariable passageways with the periphery of said rotor, each of saidpassage cooperable with a different one of said nozzles.

3. The combination as set forth in claim 2 in which a solid propellantis contained in said rotor and a plurality of circumferentially disposedspin orifices are provided through said rotor for rotating it at apredetermined rate responsive to ignition of said propellant, saidorifices communicating with said plenum chamber for providing a sourceof positive pressure thereto.

4. The combination as set forth in claim 1 in which said nozzles includea passage of a selected cross-sectional configuration to provide adesired fiow rate.

5. The combination as set forth in claim 4 in which said passage is of arectangular cross section to provide linearity of operation.

6. The combination a set forth in claim 2 in which said rotor includes aplurality of notched portions cir cumferentially extending about itsperiphery, and a nozzle is connected to said source and directed towardsaid notched portions to provide a tangential force in a planesubstantially normal to said spin axis for spinning said rotor at apredetermined rate.

7. The combination as set forth in claim 2 in which a plurality ofbattle plates are mounted radially about the inner surface of saidplenum chamber intermedia said nozzles.

8. A directional control system for maintaining a mis sile on a launchpredetermined path comprising a two degree of freedom gyroscope mountedwith its spin axis and rotor aligned with the longitudinal axis of themissile at launch, said missile movable transaxially relative to saidrotor at least three equally spaced and radially aligned jet nozzlesmounted on the missile, a pressurized plenum chamber fixed to themissile and enclosing the rotor of said gyroscope. said rotor having aperipheral grooved portion normally concentric with each of said nozzlesand providing a variable flow to said nozzles from said chamberresponsive to displacement of the missile relative to said rotor.

9. The combination as set forth in claim 8 in which said nozzlescomprise a pair of oppositely directed nozzles in the pitch plane of themissile and a pair of oppositely disposed nozzles in the yaw plane ofthe missile.

10. The combination as set forth in claim 8 in which said rotor includesa plurality of notched portions circumferentially extending in a ringabout its periphery and a spin nozzle is connected to a source ofpressurized fluid and oriented toward said notched portion for rotatingsaid rotor at a predetermined rate.

11. The combination as set forth in claim 10 in which said notchedportions are formed on the periphery of said rotor within said groovedportion and said spin nozzle extends through said plenum chamber fromsaid source of pressurized fluid.

12. A directional control system for maintaining a missile on a launchpredetermined path comprising a two degree of freedom gyroscope mountedwith its spin axis aligned with the longitudinal axis cf the missile atlaunch, at least three equally spaced and radially aligned jet nozzlesmounted on the missile. a pressurized plenum chamber fixed to themissile and enclosing the rotor of said gyroscope, said rotor having aperipheral groove portion alignable with each of said nozzles andproviding a variable flow to said nozzles from said chamber respon siveto displacement of the missile from its predetermined path, said rotorcontaining a solid propellant ignitable prior to launch and furtherincluding a plurality of spin orifices for providing rotation of saidrotor at a predetermined rate, said spin orifices disposed in a pair ofspaced concentric rings through the periphery of said rotor, each ofsaid rings proximate a different end of said plenum chamber forproviding a uniform pressurized state of said plenum chamber.

13. The combination as set forth in claim 12 in which a plurality oflongitudinally aligned battle plates are radially mounted about theinner surface of said plenum chamber in uniformly spaced groupintermediate said nozzles.

14. The combination as set forth in claim 12 in which said rotor ismounted for rotation about the stator of said gyroscope on a ballbearing raceway contained therebetween.

15. A directional control system for maintaining a missile on a launchpredetermined path comprising a two degree of freedom gyro mounted withits spin axis and rotor aligned with the longitudinal axis of themissile at launch, said missile movable transaxially relative to saidrotor, at least three substantially equally spaced and radially alignedjet nozzles mounted on the missile, a pressurized plenum chamber fixedto the missile and enclosing said rotor of said gyro, said rotordefining a plurality of passages. each of said passages normally concentric with a different one of said nozzles.

16. The combination as set forth in claim 15 in which said rotor of saidgyro comprises a spherical surface portion, said portion mounted on acorresponding spherical surface portion of the gyrostator by a pluralityof spherical bearing members.

17. The combination as set forth in claim 15 in which a common pressuresource is operative connected to said plenum chamber for pressurizing itand in which a jet connected to said source is directed at said rotor toprovide a tangential rotative force in a direction lying in a planenormal to its axis.

18. In a directional control system for a missile including a portvarying system for providing balancing forces responsive to deviation ofthe missile from a pre determined path, a gyroscope having its statorfixed to said missile and its rotor cooperable with a plurality oflateral control ports, a spherical surface portion of said stator, acorresponding spherical surface portion of said rotor, and a pluralityof spherical bearing members fixed therebetween for mounting said rotoron said stator.

References Cited by the Examiner UNITED STATES PATENTS 1,316,033 9/1919Hayden 102-50 1,316,363 9/1919 Hayden 1025O 2,703,960 3/1955 Prentiss10249 2,822,755 2/1958 Edwards et a1 10249 2,981,061 4/1961 Lilligren102-50 X 2,995,894 8/1961 Baxter et al 102-50 X FOREIGN PATENTS 646,0397/ 1962 Canada.

BENJAMIN A. BORCHELT, Primary Examiner.

V. R. PENDEGRASS, Arsixmnt Examiner.

1. A DIRECTIONAL CONTROL SYSTEM FOR MAINTAINING A MISSILE ON A LAUNCHPREDETERMINED PATH COMPRISING A GYROSCOPE MOUNTED WITH ITS SPIN AXIS ANDROTOR ALIGNED WITH THE LONGITUDINAL AXIS OF THE MISSILE AT LAUNCH, SAIDMISSILE MOVABLE TRANSAXIALLY RELATIVE TO SAID ROTOR, AT LEAST THREESUBSTANTIALLY EQUALLY SPACED AND RADIALLY ALIGNED JET NOZZLES MOUNTED ONTHE MISSILE, A SOURCE OF PRESSURIZED FLUID OPERATIVELY CONNECTIBLE TOSAID NOZZLES, SAID GYROSCOPE HAVING SAID ROTOR OPERATIVELY CONNECTEDBETWEEN SAID SOURCE AND SAID NOZZLES FOR PROVIDING A VARIABLE FLUID FLOWTHERETHROUGH TO PROVIDE BALANCING FORCES TO THE MISSILE RESPONSIVE TOITS MOVEMENT RELATIVE TO SAID ROTOR.